Bleeding air from gas turbine engine compressors is well known. Air bled from the compressor can sometimes be used to provide a source of pressurized and/or cool air to the engine or the aircraft, however air is principally bled from the compressor in order to improve the operating envelope and overall compressor efficiency, which is often expressed as improved surge margin. Increased incidence angle between the airflow and the blade leading edges at “off design” conditions tends to cause separation of the flow on the suction side of the blades, which results in blade stall and eventually complete surging of the compressor. By bleeding off this stalled airflow adjacent the blade tips, the surge margin of the compressor is thus increased. This accordingly improves the overall efficiency of the compressor.
However, separation of airflow on the compressor blades can also result from factors other than increased blade leading edge incidence. Particularly, the interaction between the boundary layer formed on a stationary outer shroud and a shock wave produced by supersonic compressor blade tips rotating within the shroud, also tends to cause additional flow separation which can result in blade stall and to full compressor surge. Although the inlet flow may be subsonic in a subsonic compressor, the flow relative to the rotor blade tips of a high speed compressor can nevertheless become supersonic, causing separation-inducing shock waves at the blade tips.
Accordingly, there is a need to provide an improved compressor which addresses these and other limitations of the prior art, and it is therefore an object of this invention to do so.